1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to a film cooling hole for a turbine airfoil.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
Turbine airfoils (which include rotor blades and stator vanes) include film cooling holes to discharge a layer of film cooling air over a surface of the airfoil to form a blanket of cool air against the hot gas stream that flows over the surface. In one prior art film cooling hole, the film hole passes straight through the airfoil wall at a constant diameter and exits at an angle to the surface. Some of the cooling air is ejected directly into the mainstream gas flow causing turbulence, coolant dilution and a loss of downstream film effectiveness. Also, the hole breakout in the streamwise elliptical shape will induce stress in a blade application. FIGS. 1 through 7 shows varies views of the straight film cooling hole with constant diameter.
FIGS. 8 through 10 show another prior art film cooling that includes a diffusion section. This film hole includes a 10×10×10 streamwise three dimension diffusion hole. This film hole includes a constant cross section flow area at an inlet end for metering the cooling air flow and a diffusion section downstream. The diffusion section includes three walls each having 10 degrees of slant. The upstream wall of the film hole (the left side in FIG. 9) has zero diffusion and is parallel to the film hole axis. In this film hole, hot gas from the mainstream flow frequently gets entrained into the upper corner and causes shear mixing with the cooling air (see FIG. 11 reference numeral 12). This results in a reduction of film cooling effectiveness for the film cooling hole. Also, internal flow separation occurs (FIG. 11 reference numeral 11) within the diffusion hole at a junction between the constant cross section area and the diffusion region.